The present invention relates generally to spacecraft attitude pointing methods, and more particularly, to spacecraft attitude pointing methods that provide for stepping multiple appendages to reduce spacecraft attitude pointing disturbances caused by appendage stepping and cancellation of solar array slew disturbances.
A class of spacecraft known as three-axis stabilized spacecraft employ a solar array to generate power for the spacecraft. The solar array must be maintained in a position normal to the sun to absorb the optimum amount of radiation. Because the solar array is maintained normal to the sun, a servo controlled stepping mechanism, such as a stepping motor and an appropriate gear train, is typically employed to cause the solar array to track the sun while the spacecraft is in constant rotation relative to the sun in an orbit about the earth. Other types of attitude control mechanisms, such as dc motors, prove to be relatively difficult to control and are heavy. However, in theory, servo controlled dc motors would not generate oscillation. It is desirable to use stepper motors because stepper motors are relatively simple to control, reliable, lightweight and well adapted to continuous use.
One of the major problems with the use of stepping motors is that the stepping action can excite a highly flexible array such that oscillation is induced within the spacecraft. The induced oscillation is particularly critical in spacecraft where absolute platform stability is desirable or required, such as platforms for high resolution optical imaging systems. Vibrations can cause deterioration of any inertia-sensitive operations of a spacecraft. Therefore, it is desirable to solve the problem of induced oscillation caused by a stepper motor.
U.S. Pat. No. 4,843,294 entitled "Solar Array Stepping to Minimize Array Excitation" assigned to the assignee of the present invention discloses one way to improve spacecraft attitude pointing. The method disclosed in U.S. Pat. No. 4,843,294 deadbeats individual appendage oscillations. As such, the stepping of the solar array wings were stepped in a manner that minimized their individual oscillations. The present invention improves upon the teachings of U.S. Pat. No. 4,843,294.
More particularly, and in accordance with the teachings of U.S. Pat. No. 4,843,294, mechanical oscillations of a mechanism containing a stepper motor, such as a solar-array powered spacecraft, are reduced and minimized by the execution of step movements in pairs of steps. The period between steps is equal to one-half of the period of torsional oscillation of the mechanism. Each pair of steps is repeated at needed intervals to maintain desired continuous movement of the portion of elements to be moved, such as the solar array of a spacecraft. In order to account for uncertainty as well as slow change in the period of torsional oscillation, a command unit may be provided for varying the interval between steps in a pair.
Furthermore, solar arrays are Sun tracking, while satellite payloads are Earth tracking. This means the solar arrays rotate with respect to the body of the spacecraft. Every step in the rotation causes a disturbance. As solar arrays become physically larger, so do the disturbances caused by rotation the solar arrays. Previous systems developed by the assignee of the present invention relied entirely on feedback to reduce the disturbances. The present invention takes apriori knowledge of an event (solar array step) and uses that knowledge to reduce the disturbance.
Accordingly, it is an objective of the present invention to provide for spacecraft attitude pointing methods that provide for stepping multiple appendages to reduce spacecraft attitude pointing disturbances caused by appendage stepping and cancellation of solar array slew disturbances.